Gas turbine engines are internal combustion engines typically used to provide thrust to an aircraft or to provide power for land-based applications. In general, a gas turbine engine may consist of a fan section and a core engine located downstream of the fan section. In an upstream to downstream direction, the core engine may generally include: 1) a compressor section, which may include a low pressure compressor (LPC) and a high pressure compressor (HPC) located downstream of the LPC, 2) one or more combustors, and 3) a turbine section, which may include a high-pressure turbine (HPT) and a low-pressure turbine (LPT) located downstream of the HPT. During operation, a portion of the air drawn into the fan section may be routed through a primary flowpath defined by the core engine. In the core engine, the air may be pressurized in the compressor section and then mixed with fuel and combusted in the combustor(s) to generate hot combustion gases. The hot combustion gases may then expand through and drive the turbine section which extracts energy from the hot combustion gases to power the compressor section and the fan section.
Certain components of the gas turbine engine, such as the rotating blades of the turbine section, may become hot during operation. As a result, many gas turbine engines incorporate a cooling system that supplies cooling air to the turbine section or other heat-susceptible structures. For example, in legacy gas turbine engine designs, a portion of the primary airflow exiting the high pressure compressor may be diverted through a path located radially inboard of the combustor to provide a ‘turbine cooling air’ to the HPT. However, this approach may reach its limitation as the cooling air reaches temperatures above the temperature limit of the material properties forming the turbine components.
A more advanced variation, as disclosed in U.S. Patent Application Publication Number 2013/0219917, cools the turbine cooling air by mixing it with a heat-exchange cooled ‘buffer air’ diverted from a bearing compartment to generate a ‘buffer cooled cooling air’(or BCCA) that is then delivered to the HPT for cooling the turbine blades as well as other turbine structures. More specifically, this approach involves diverting a fraction of the turbine cooling air through a heat exchanger to generate the lower temperature buffer air that is then routed through the bearing compartment to cool the structures of the compartment as well as to buffer the bearing compartment seals. A fraction of the buffer air flowing through the bearing compartment is then diverted from the bearing compartment cooling air path through multiple tubes to join with the turbine cooling air to generate the BCCA that is then supplied to the HPT. While effective, a significant disadvantage of this approach is that it may require stronger/heavier bearing compartment structures and passageways that are capable of withstanding greater air volumes and pressures caused by the additional buffer air used to cool the turbine cooling air. In addition, the higher air pressures in the bearing compartment may present challenges for establishing acceptable pressure differentials across the carbon seals of the bearing compartment. Even further, the multiple tubes that carry the buffer air away from the bearing compartment to mix with the cooling air may create a difficult assembly process and add even further weight to the cooling system.
Clearly, there is a need for improved strategies for providing cooling air to turbine sections of gas turbine engines.